The present invention relates to a system for protecting structures or components adjacent to ion accelerators from erosion or other damage by impact of accelerated ions with such structures and/or by impact by sputtered atoms resulting from the accelerated ions contacting other surfaces.
Electric propulsion thrusters use steady, quasi-steady, oscillating and transient electric and magnetic fields to accelerate ions to high exhaust velocity. Due to their high exhaust velocity, electric thrusters are finding increased application on commercial and government satellites. Significant propellant mass savings is realized by increasing the thruster exhaust velocity. Electric thrusters typically accelerate ions to exhaust velocities which are much greater ( greater than 15,000 m/sec) than present-day hydrazine monopropellant engines (2200 m/sec). A critical issue associated with electric thruster implementation is the erosion and sputtering which occurs when high-energy ions impinge upon spacecraft surfaces.
For example, two high-energy thrusters in current use today are the Hall thruster and ion engine. Both thrusters use xenon as a propellant and emit high-energy ions axisymmetrically in the 2xcfx80 space from the exit plane. Thruster plumes are typically differentiated from each other by the mean ion energy, and by how collimated the ion beam is as characterized by the xe2x80x9cplume divergence anglexe2x80x9d. The plume divergence angle is the half angle, from the thruster centerline, which encloses 95% of emitted ions. The Hall thruster typically has a plume divergence angle of 40-45 degrees, with xenon ion energies in the 200-400 eV range. The ion engine is much more collimated with a plume divergence angle of approximately 20-25 degrees and ion energies greater than 1000 eV. The plasma created by both thrusters is quasineutral throughout the plume.
Many spacecraft configurations, in particular those of geosynchronous satellites, require that some external spacecraft components (e.g. spacecraft solar arrays, antennas and thermal heat rejection panels) be positioned within the electric thruster plume. One integration issue to be resolved is the sputter erosion of these components. The obvious effect of such sputtering is the degradation of the adjacent component. In the case of a solar panel, the undesirable result can be a decrease in the solar array power generation; in the case of antennas, antenna performance reductions may occur; and in the case of thermal heat rejection panels, the problem could be a loss of thermal control.
Another issue is the redeposition of sputtered atoms (released as a result of impact by high energy ions in the plume) onto sensitive spacecraft surfaces such as optics and thermal heat rejection panels. The deposition rates associated with low-energy sputtered atoms and ions are difficult to predict for a number of reasons. First, the distribution of sputtered atom ejection velocities and charge state is usually unknown. Second, their motion is largely influenced by the local electric fields near spacecraft surfaces, which are difficult to predict. Third, sputtered atoms will interact with the high energy ions from the thruster (via charge exchange reaction) to create positively charged metal ions, which are subsequently forced by electric fields in the plume toward spacecraft surfaces.
In general, the above issues are engineered using a number of methods. One method, used on geosynchronous satellites, is to cant the thruster away from sensitive surfaces, in particular the solar arrays. Cant angles are typically on the order of 45 degrees from the optimum thrusting direction, which results in a significant thrusting efficiency and propellant loss. Due to canting, thruster lifetime requirements are increased, resulting in expensive qualification life tests extending up to 10,000 hours. In addition to canting, degradation can be minimized by optimizing the position of critical surfaces with respect to the thruster and by adding additional margin in surface thicknesses, power budgets, etc. Such optimization is performed under the constraints associated with providing the optimum spacecraft configuration. This approach requires a detailed model of the motion of high-energy ions and deposition products in conjunction with validation testing and fundamental measurements of sputtering processes. Modeling and testing associated with this approach are time consuming, and carry the risk of not modeling all of the relevant physics. Another approach to mitigating contamination issues is to limit thruster operation to those times when the external component is within an acceptable position relative to the plume. The thruster firing keep-out zone approach has the drawback of complicating mission operations, thus increasing the total mission cost.
All of the above mitigation approaches will become increasingly difficult to engineer in the future for the following reasons. First, the trend in satellite capability is toward higher power levels and longer thruster firing times. These trends increase the total exposure of high-energy ions impinging onto spacecraft surfaces. Thrusters in operation currently operate at power levels of 0.5 to 4.0 kW. Future missions have projected power levels as high as 50 kW. The second trend in spacecraft design is toward the placement of more components outside of the spacecraft and further out from the spacecraft. Examples include larger and more antennas, larger solar arrays, larger radiation panels, all of which will make the spacecraft system design task increasingly more difficult with respect to integrating electric thrusters.
Yet another approach has been to add a flat plate called a plume shield between the thruster and a sensitive surface or component. Typically a plume shield is formed of low-sputter yield material. The shield screens spacecraft components from high-energy ions. However a drawback associated with known plume shields is that the sputter products from the plume shield itself will contaminate sensitive spacecraft surfaces. As stated earlier, there are many mechanisms leading to the contamination of critical surfaces. Thus incorporating known plume shields still requires detailed trade studies, analysis, and tests.
An example of a plume shield is described in a publication by S. Shimada et al., xe2x80x9cDevelopment of Ion Engine System for ETS-VI,xe2x80x9d Proceedings of the 23rd International Propulsion Converence, Seattle, Wash., U.S.A., September 1993, Paper No. IEPC 93-009, vol. 1, pp. 116-124. The plume shield described in this publication was mounted very close to the thruster and was constructed of an aluminum alloy in the known flat plate configuration. The thruster power level was low, only 700 W. The primary purpose of the plume shield was to protect the spacecraft thermal control surfaces from the deposition of molybdenum eroded from the ion engine grids.
A plume shield in accordance with the present invention significantly reduces the creation of sputter products, yet still performs the function of screening high-energy ions. The benefits of a low-sputter product plume shield are significant. It allows for possible reduction of the cant angle, which is equivalent to a thruster performance increase. It mitigates the need for thruster on-time keep-out zones, and thus allows for significant cost savings in mission operations and mission design. It allows more design freedom in positioning thrusters and spacecraft external components, and thus significantly reduces the cost associated with integrating electric thrusters on spacecraft. From a plume impacts point of view, this invention enables the use of very high power ( greater than 50 kW) electric thrusters on geosynchronous spacecraft.
In one embodiment, a plume shield in accordance with the present invention incorporates a series of long slats, preferably parallel and spaced apart uniformly, extending from a base plate. This xe2x80x9clouveredxe2x80x9d design forms a series of cavities between the slats which trap high energy ions and atoms sputtered from the shield. In a preferred design, the slats extend lengthwise in a direction generally parallel to the axis of the plume, and have their widths extending radially. The outer marginal portions of the slats can be angled to form a target area or segment which can assist in trapping sputtered atoms and preventing them from escaping from the shield.
In another aspect of the present invention, the effectiveness of the shield in trapping and retarding ions and sputtered atoms is accomplished by biasing the potential of the shield with respect to the spacecraft structure. This can result in slowing ions to energies below the sputter threshold energy. Magnets may be used to create a magnetic field parallel to a target area of the shield, which can limit electron current to the target.
Other aspects of the invention are described in more detail below.